This invention relates to attitude control of spacecraft, and more particularly to unloading of reaction or momentum wheel momentum by the use of controlled magnetic torquers.
Orbiting spacecraft are used for a large variety of sensing and communication purposes for photographic purposes, it is desirable for the spacecraft to be relatively near the Earth, so that the cameras or sensors are close to the subjects, and so that the spacecraft may cover large amounts of surface by virtue of a rapid orbital rate. For communication purposes, a geosynchronous equatorial orbit is often desirable. Whatever the orbit, the satellite must be stabilized in space if the sensors or antennas are to be pointed in appropriate direction.
Spacecraft attitude stabilization may be accomplished by spinning the spacecraft and by mounting the sensors or antennas on a despun platform. Alternatively, the spacecraft may be stabilized in three axes. Three-axis stabilization may be accomplished by a control system using fuel-burning thrusters, but the use of such thrusters requires the expenditure of fuel, which tends to limit the service life of the spacecraft. Another method of three-axis stabilization uses reaction wheels or momentum wheels mounted within the spacecraft. The reaction wheels are torqued under the control of an attitude control system to cause the spacecraft body to assume the desired attitude. The energy for torquing the wheels is electrical, and is derived from solar panels. Attitude control by the use of wheels therefore generally provides longer operating lifetime than simple attitude control by the use of thrusters.
In order to maintain a face of the satellite directed toward the Earth, the satellite must rotate about its pitch axis at the orbit rate. When using reaction wheel control, the wheel or wheels are torqued to maintain an earth-oriented attitude. Environmental disturbance torques that act on the spacecraft tend to result in a gradual increase in the wheel momenta. In order to prevent the wheels from reaching their mechanical momentum limits, and to thereby maintain the ability to provide reaction wheel based attitude control, the wheel momentum must be periodically reduced. Wheel momentum may be reduced by the use of fuel-consuming thrusters which torque the spacecraft body in a manner which opposes the wheel momentum. However, the use of such thrusters is subject to the same objection of limited lifetime described above.
FIG. 1a illustrates a spacecraft designated generally as 10, and including a body 12 and a solar panel 14. Spacecraft 10 is following an orbit illustrated by dashed-line path 16 about the Earth 18. The Earth's rotational axis is illustrated as 20, and the Equator is illustrated by dash-line 22. Lines of magnetic force associated with the Earth's magnetic field are illustrated as 24. As illustrated in FIG. 1, solar panel 14 is directed toward the sun 26, and is controlled to continue facing the sun at all orbital positions of satellite 10.
FIG. 1b illustrates details of spacecraft 10. In FIG. 1b a first reaction wheel 30 is oriented with its axis along the pitch or y axis, a second reaction wheel 32 is oriented with its axis along the x or roll axis, and a third reaction wheel 34 is oriented with its axis along the yaw or z axis. Also in FIG. 1b, magnetic torquer windings illustrated as 36, 38 and 40 are wound about the spacecraft body. Further in FIG. 1b, a set of three-axis magnetometers is illustrated as a block 41, and is connected to a torquer control arrangement illustrated as a block 42
FIGS. 2a, 2b and 2c illustrate as time functions the roll, pitch and yaw disturbance torques, respectively, measured in inch-lb. In FIG. 2, the disturbance torques include a periodic component and an offset component. The disturbance torques arise from atmospheric drag, gravity gradient effects, and solar pressure. The cyclical perturbations recur at roughly 6,000 second intervals, corresponding to the orbital period. The constant offset of the disturbance torques is attributable to the asymmetric orientation of solar panel 14 of FIG. 1a. In FIG. 2, zero seconds and each 6,000 second increment thereafter represent the times at which the spacecraft crosses the Earth's Equator, known as the ascending node.
FIGS. 3a, 3b and 3c represent as time functions the roll, pitch and yaw wheel momentum, respectively, in the absence of unloading. The changes in wheel momentum with time are attributable to the disturbance torques illustrated in FIG. 2. Some components of the disturbance torques of FIG. 2 contribute to the cyclical variations of wheel momentum in FIG. 3, while others result in secular or continually increasing or decreasing components. For example, in FIG. 3a, the peak values of the cyclical component of wheel momentum increases with increasing time. At some point, therefore, such a continuing increase in amplitude will result in reaching the wheel momentum limit. Similarly, the pitch wheel momentum of FIG. 3b increases monotonically with increasing time. In order to maintain continuous attitude control, the secular component of the wheel momentum must be unloaded.
FIGS. 4a, 4b and 4c represent the roll, pitch and yaw torquer magnetic moment measured in ampere-turns-meter.sup.2 (ATM.sup.2) for a spacecraft using a prior art torquer control system which pulse width or continuously modulates the torquer current. The prior art control system measures the Earth's geomagnetic field B and the net momentum of the reaction wheel, and forms a vector cross product HXB. The normalized HXB vector is multiplied by the maximum magnetic moment which can be achieved along each of the roll, pitch and yaw axes to determine the torquer dipole command. In FIGS. 4a, 4b and 4c, large excursions or jitter in the roll and pitch moments, and lesser excursions in the yaw moments, occur at about 7500 seconds, 14,000 seconds, and 20,000 seconds, corresponding to passage of the spacecraft across the Earth's magnetic poles. The jitter illustrated in FIG. 4 arises because the magnitude of the Earth's geomagnetic field near the poles is large, but the cyclical component of the wheel momentum results in relatively low wheel momentum. Operation of the magnetic torquers in regions near the poles, therefore, produces a high magnitude of wheel momentum change at times when the wheel momentum is low, which tends to drive the torquer command from one extreme to the other. This in turn results in the illustrated jitter, which disturbs the spacecraft attitude. Such disturbances are undesirable.